CN110206647B

CN110206647B - Aeroengine bearing support subassembly and aeroengine

Info

Publication number: CN110206647B
Application number: CN201810167754.4A
Authority: CN
Inventors: 赵芝梅; 唐振南; 宋会英; 郑李鹏
Original assignee: AECC Commercial Aircraft Engine Co Ltd
Current assignee: AECC Commercial Aircraft Engine Co Ltd
Priority date: 2018-02-28
Filing date: 2018-02-28
Publication date: 2020-08-18
Anticipated expiration: 2038-02-28
Also published as: CN110206647A

Abstract

The invention aims to provide an aeroengine bearing support assembly and an aeroengine. The aeroengine bearing support component comprises a mounting seat, a rigid support structure and a shock absorption component; the shock absorption component comprises a first connecting ring, a second connecting ring and a shock absorption body made of damping energy-absorbing materials; the first connecting ring and the second connecting ring are respectively and fixedly connected to the mounting seat in a sealing manner and are respectively and hermetically connected with the bearing so as to support the bearing, and a closed annular cavity is formed between the mounting seat and the bearing; the damping energy-absorbing material is used for filling the annular cavity to form the shock-absorbing body. Because the existence of inclosed annular cavity, inhale the shake the body and can hold by inclosed annular cavity all the time bearing the impact of each direction and the in-process that warp repeatedly, and then make inhale the shake ability that shakes the body and can keep for a long time to it has longer live time to make to inhale the shake subassembly. An aircraft engine includes the aircraft engine bearing support assembly described above and therefore has a shock absorbing assembly with a longer service life.

Description

Aeroengine bearing support subassembly and aeroengine

Technical Field

The invention relates to an aeroengine bearing assembly and an aeroengine.

Background

According to the requirements of airworthiness regulations (FAR33.74, FAR33.94), commercial aircraft engines must ensure that the occurrence of an FBO event (fan blade fly-out) does not result in catastrophic consequences. The FBO event can generate a very large unbalanced load, which may cause damage to key components such as an engine mount joint and a bearing, and cause catastrophic accidents such as engine falling, blade breakage and cabin breakdown. Immediately after the FBO event, the damaged engine should be stopped to slowly bring the engine down from a higher operating speed to windmill rotation and continue for a period of time (sometimes up to 180 minutes) during the windmill rotation phase until deceleration for safe landing is achieved when landing conditions are met.

In order to ensure that the engine can land safely after an FBO event occurs, a load reduction design is generally adopted to reduce the FBO limit load borne by each component (mainly including a mounting system, a low-pressure rotating shaft and the like) of the engine, so that the safety of the engine is ensured. A common load relief design, also called a fusing design, is to provide a component with weak mechanical properties, such as a thinning section (see patent US6447248), a reducing bolt (see patent US7318685) and the like, on a bearing support structure, so that the component fails under the action of a predetermined load (threshold value), on one hand, the critical rotation speed of the rotor is reduced, the rotor is in a supercritical rotation speed state, and an unbalanced load is reduced; on the other hand, the transmission path of the FBO load to the stator casing is changed, so that the FBO load is redistributed, and the safety of the engine is effectively protected. The current common fusing design is generally located on the No. 1 bearing seat or the supporting cone wall, and under the FBO load, the local weak part is broken to change the transmission path, reduce the supporting rigidity and reduce the FBO load transmitted to the key part, but the following problems still exist:

1. after the fusing component fails, the No. 1 bearing does not have any supporting effect on a fan shaft, and the low-pressure rotor completely loses the constraint at the No. 1 bearing, so that the fan swings greatly, the No. 1 bearing lubricating and cooling system possibly fails, and the bearing fails. In addition, excessive fan oscillation can produce a large unbalanced moment on the adjacent bearing # 2. This can result in large local bending deformations and stress concentrations of the fan shaft at bearing number 2, which can cause the fan shaft to break and the fan rotor to fly off, with catastrophic consequences. 2. For the fusing design scheme that a local weak part (such as patent US6447248) is adopted on the bearing supporting conical wall No. 1, after the local weak part is broken and fails under the action of FBO load, the front half section of the supporting conical wall is completely out of restraint, can move randomly around a fan shaft, and generates friction and scratch with the fan shaft, so that the fan shaft is damaged. Since the engine also needs to last long (perhaps as long as 180 minutes) during the windmilling period after the FBO event occurs, damage to the fan shaft may affect the safety of the engine under continuous rotation.

As shown in fig. 4, US20160097301a1 discloses a shock absorbing assembly for an aircraft engine dealing with FBO events, comprising a connection 40 to a bearing outer race 39, the connection 40 being provided with a pre-breaking point 33, and a shock absorbing member 34. When an FBO event occurs, the pre-break point 33 of the connector 40 breaks such that the shock absorbing member 34 provides a cushion radially between the fan bearing structure 37 and the bearing outer race 39. Since the engine also needs to last a long time (perhaps as long as 180 minutes) during the fan spin phase after the FBO event occurs, this places high demands on the time that shock absorbing member 34 can last in use.

There is a need in the art for an aircraft engine bearing support assembly and aircraft engine that provides a shock absorbing assembly having a longer useful life after an FBO event has occurred.

Disclosure of Invention

It is an object of the present invention to provide an aircraft engine bearing support assembly having a shock absorbing member with a longer service life.

It is also an object of the present invention to provide an aircraft engine including an aircraft engine bearing support assembly as described above, which therefore has a shock absorbing assembly with a longer service life.

The aeroengine bearing support assembly for achieving the purpose is used for supporting a bearing and comprises a mounting seat, a rigid support structure and a shock absorption assembly, wherein the rigid support structure is fixedly connected to the mounting seat and is used for being connected with the bearing to support the bearing;

the shock absorption component comprises a first connecting ring, a second connecting ring and a shock absorption body made of damping energy-absorbing materials;

the first connecting ring and the second connecting ring are respectively and fixedly connected to the mounting seat in a sealing manner, are used for being respectively and hermetically connected with the bearing, and jointly define a closed annular cavity together with the mounting seat and the outer ring of the bearing; the damping energy-absorbing material is used for filling the annular cavity to form the shock-absorbing body;

the first connecting ring, the second connecting ring and the shock absorbing body are used for bearing the load between the mounting seat and the bearing to generate deformation, so that impact energy is absorbed.

The aeroengine bearing support assembly is further characterized in that the aeroengine further comprises an intermediate casing and a fan shaft, the bearing comprises an inner ring, a roller and an outer ring, and the roller is clamped between the inner ring and the outer ring;

the inner ring is fixedly connected with the fan shaft, and the mounting seat is fixedly connected to the intermediate casing; the mounting seat surrounds the outer ring and shares a central axis with the bearing.

The aero-engine bearing support assembly is further characterized in that the first connecting ring and the second connecting ring are annular plates centered on the central axis and axially spaced to form the annular cavity; the annular cavity is centered on the central axis.

The aircraft engine bearing support assembly is further characterized in that the bearing is a number 1 bearing of the aircraft engine.

The aircraft engine bearing support assembly is further characterized in that, with radial reference, the inner edge of the first connecting ring is in sealing connection with the outer ring, and the outer edge of the first connecting ring is in sealing connection with the mounting seat; and the inner edge of the second connecting ring is in sealing connection with the outer ring, and the outer edge of the second connecting ring is in sealing connection with the mounting seat.

An aircraft engine for achieving the object, comprising a bearing, the aircraft engine further comprising an aircraft engine bearing support assembly as claimed above, the aircraft engine bearing support assembly comprising a mount, a rigid support structure and a shock absorbing assembly, the rigid support structure being fixedly attached to the mount and being connected to the bearing for supporting the bearing;

the first connecting ring and the second connecting ring are respectively and fixedly connected to the mounting seat in a sealing manner, are respectively and hermetically connected with the bearing so as to support the bearing, and jointly define a closed annular cavity with the mounting seat and the outer ring of the bearing; the damping energy-absorbing material is filled in the annular cavity to form the shock-absorbing body;

the rigid support structure and the shock absorbing member together support the outer race of the bearing before failure of the rigid support structure, and the shock absorbing member alone supports the outer race after failure of the rigid support structure, and the primary connecting ring, the secondary connecting ring, and the shock absorbing body are adapted to receive a load between the mount and the bearing and to deform, thereby absorbing impact energy.

The aircraft engine is further characterized by further comprising an intermediate casing and a fan shaft, wherein the bearing comprises an inner ring, a roller and an outer ring, and the roller is clamped between the inner ring and the outer ring;

the inner ring is fixedly connected with the fan shaft, and the mounting seat is fixedly connected to the intermediate casing; the mounting seat surrounds the outer ring and shares a central axis with the bearing, the rigid supporting structure is connected with the outer ring, and the first connecting ring and the second connecting ring are respectively connected with the outer ring in a sealing manner.

The aircraft engine is further characterized in that the first connecting ring and the second connecting ring are both annular plates taking the central axis as the center and are arranged at intervals along the axial direction to form an annular cavity; the annular cavity is centered on the central axis.

The aircraft engine is further characterized in that the bearing is a No. 1 bearing of the aircraft engine.

The aircraft engine is further characterized in that the inner edge of the first connecting ring is in sealing connection with the outer ring and the outer edge of the first connecting ring is in sealing connection with the mounting seat by taking the radial direction as a reference; and the inner edge of the second connecting ring is in sealing connection with the outer ring, and the outer edge of the second connecting ring is in sealing connection with the mounting seat.

The positive progress effects of the invention are as follows: the invention provides an aeroengine bearing support assembly, which is used for supporting a bearing and comprises a mounting seat, a rigid support structure and a shock absorption assembly, wherein the rigid support structure is fixedly connected to the mounting seat and is used for being connected with the bearing so as to support the bearing; the shock absorption component comprises a first connecting ring, a second connecting ring and a shock absorption body made of damping energy-absorbing materials; the first connecting ring and the second connecting ring are respectively and fixedly connected to the mounting seat in a sealing manner and are respectively and hermetically connected with the bearing so as to support the bearing, and a closed annular cavity is formed between the mounting seat and the bearing; the damping energy-absorbing material is used for filling the annular cavity to form a shock-absorbing body; the primary connecting ring, the secondary connecting ring and the shock absorbing body are used for bearing the load between the mounting seat and the bearing to generate deformation, so that the shock energy is absorbed.

Because the existence of inclosed annular cavity, inhale the shake the body and can hold by inclosed annular cavity all the time bearing the impact of each direction and the in-process that warp repeatedly, therefore the damping energy-absorbing material that constitutes to inhale the shake body can not disperse and drop, but remains throughout in inclosed annular cavity, and then makes the shock-absorbing capacity who inhales the shake body can keep for a long time to it has longer live time to make to inhale the shake subassembly.

Drawings

The above and other features, properties and advantages of the present invention will become more apparent from the following description of the embodiments with reference to the accompanying drawings, in which:

FIG. 1 is a cross-sectional view of an aircraft engine according to the present invention;

FIG. 2 is an enlarged view taken at B of FIG. 1 showing the bearing and an aircraft engine bearing support assembly;

FIG. 3 is an enlarged view at C of FIG. 2;

figure 4 is a schematic view of a shock absorbing member of an aircraft engine of the prior art.

Detailed Description

The present invention is further described in the following description with reference to specific embodiments and the accompanying drawings, wherein the details are set forth in order to provide a thorough understanding of the present invention, but it is apparent that the present invention can be embodied in many other forms different from those described herein, and it will be readily appreciated by those skilled in the art that the present invention can be implemented in many different forms without departing from the spirit and scope of the invention.

In the description of the present invention, it is to be understood that the terms "length", "width", "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner" and "outer" etc. indicate orientations or positional relationships based on those shown in the drawings, and are only for convenience of description and simplicity of description, and do not indicate or imply that the referenced devices or elements must have a particular orientation, be constructed and operated in a particular orientation, and thus, are not to be construed as limiting the present invention.

Furthermore, the terms "first", "second" and "first" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance or implicitly indicating the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include one or more of that feature. In the description of the present invention, "a plurality" means two or more unless specifically defined otherwise.

In the present invention, unless otherwise expressly stated or limited, the terms "mounted," "connected," "secured," and the like are to be construed broadly and can, for example, be fixedly connected, detachably connected, or integrally formed; can be mechanically or electrically connected; either directly or indirectly through intervening media, either internally or in any other relationship. The specific meanings of the above terms in the present invention can be understood by those skilled in the art according to specific situations.

It should be noted that fig. 1-4 are exemplary only, are not drawn to scale, and should not be construed as limiting the scope of the invention as actually claimed. The terms "axial," "radial," and circumferential in the present disclosure are used with reference to the orientation of aircraft engine 100.

FIG. 1 illustrates an embodiment of an aircraft engine 100 of the present invention, aircraft engine 100 having a central axis A-A, and only one half of a cross-sectional view of aircraft engine 100 is shown, with portions not shown being entirely symmetrical with portions shown about central axis A-A. The axis surrounded by the "ring" in the present invention is the central axis a-a.

With continued reference to fig. 1, aircraft engine 100 includes a fan rotor 3, an intermediate case 4, bearings 1 and aircraft engine bearing support assembly 2, a fan shaft 5 and an extension shaft 6. Bearing 1 is in the embodiment shown in fig. 1 the front bearing of an aircraft engine 100, i.e. bearing number 1. In further embodiments, the aero engine bearing support assembly 2 may also be used to support bearings in other locations, such as the aft support bearing 7, i.e., bearing No. 2.

Embodiments of the present invention will now be described in detail, taking as an example the application of an aircraft engine bearing support assembly 2 to bearing number 1.

As shown in connection with fig. 1, 2 and 3, in this embodiment, the bearing 1, i.e. bearing No. 1, comprises an inner ring 10, rollers 11 and an outer ring 12, the inner ring 10 being fixedly connected to the fan shaft 5, the outer ring 12 being fusably connected to the intermediate casing 4 via the aero-engine bearing support assembly 2.

In one embodiment, the forward end of fan shaft 5 is supported by bearing number 1, and supporting cone wall 8 connects bearing number 1 to intermediate case 4, and supporting cone wall 8 is a significant path for fan rotor loads to be transferred to the intermediate case.

In the above embodiment, the fan shaft 5 is supported by the bearing number 1 and the bearing number 2 together, wherein the bearing number 1 is a rolling rod bearing and provides radial constraint for the fan shaft 5; the bearing number 2 7 is a ball bearing that provides both axial and radial restraint to the extension shaft 6 of the fan shaft 5.

To reduce the bearing stiffness of bearing number 1 after an FBO event occurs, and to reduce the load transmitted from fan rotor 3 to intermediate case 4, the structure of aero-engine bearing assembly 2 in one embodiment of the present invention is described below.

With reference to fig. 1, 2 and 3, an aircraft engine bearing support assembly 2 comprises a mounting block 20, a rigid support structure 21 and a shock absorbing assembly 22, the rigid support structure 21 being fixedly attached to the mounting block 20 and adapted to be coupled to a bearing 1 to support the bearing 1.

In one embodiment, the rigid support structure 21 has a weakened portion 21a, and the weakened portion 21a is adapted to break upon receiving an overload force, thereby allowing the shock absorbing member 22 to solely support the bearing 1. The weakened portions 21a may be grooves discretely distributed in the circumferential direction as shown in fig. 3, or may be any other structures with locally weakened strength, and may be continuously or discontinuously distributed in the circumferential direction, as long as the condition for causing the rigid support structure 21 to fail under the FBO load is satisfied.

The rigid support structure 21 and the shock absorbing member 22 together support the outer race 12 of the bearing 1 before failure of the rigid support structure 21, and the shock absorbing member 22 alone supports the outer race 12 and provides a shock absorbing effect after failure of the rigid support structure 21.

With continued reference to fig. 1, 2 and 3, the shock absorbing member 22 includes a primary connecting ring 221, a secondary connecting ring 222 and a shock absorbing body 220 formed of a damping energy absorbing material; the first connection ring 221 and the second connection ring 222 are respectively and fixedly connected to the mounting base 20 in a sealing manner, and are used for being respectively and fixedly connected with the bearing 1 in a sealing manner so as to support the bearing 1, and define a closed annular cavity 2a together with the mounting base 20 and the outer ring 12 of the bearing 1; the damping energy-absorbing material is used for filling the annular cavity 2a to form a shock-absorbing body 220; the primary connection ring 221, the secondary connection ring 222 and the shock-absorbing body 220 are used to receive a load between the mount 20 and the bearing 1 to be deformed, thereby absorbing impact energy. In one embodiment, the first and second connection rings 221 and 222 are thin-walled metal plates, such as aluminum alloy plates. The damping energy-absorbing material is metal rubber or foamed aluminum.

Because of the existence of the closed annular cavity 2a, the shock-absorbing body 220 can be always accommodated by the closed annular cavity 2a in the process of bearing the impact in all directions and repeatedly deforming, so that the damping energy-absorbing material forming the shock-absorbing body 220 can not disperse and fall off, but always keeps in the closed annular cavity 2a, and further the shock-absorbing capacity of the shock-absorbing body 220 can be kept for a long time, so that the shock-absorbing component 22 has a long service life.

With continued reference to fig. 3, in one embodiment, the first connection ring 221 has a first curved structure 221a having a wave shape, the second connection ring 222 has a second curved structure 222a having a wave shape, and the shock absorbing body 220 is snap-fitted with the first curved structure 221a and the second curved structure 222a, respectively. The first and

second bending structures

221a and 222a can make the first and second connection rings 221 and 222 have a force direction in which deformation is relatively easy to occur.

In a more specific embodiment, the mounting seat 20 surrounds the outer ring 12 and is coaxial with the bearing 1 about a central axis a-a, the rigid support structure 21 is connected to the outer ring 12, and the first and second connection rings 221 and 222 are sealingly connected to the outer ring 12, respectively. The first connection ring 221 and the second connection ring 222 are both annular plates centered on the central axis a-a, and are arranged at intervals in the axial direction to form an annular cavity 2 a; the annular cavity 2a is centered on the central axis a-a.

In order to ensure the sealing performance of the annular cavity 2a, the inner edge of the first connecting ring 221 is connected with the outer ring 12 in a sealing manner, and the outer edge of the first connecting ring 221 is connected with the mounting seat 20 in a sealing manner; with radial reference, the inner edge of the secondary connecting ring 222 is sealingly connected to the outer ring 12, and the outer edge of the secondary connecting ring 222 is sealingly connected to the mounting seat 20.

After an FBO event occurs, the fan rotor 3 generates a large impact and unbalance load and the predetermined weak portion 21a of the rigid support structure 21 fails. In this case, only the shock absorbing member 22 is connected between the outer race 12 of the bearing 1 and the mount 20. Under the FBO load, the shock-absorbing element 22 deforms according to a predetermined deformation, absorbing on the one hand the energy transmitted from the bearing 1 and the supporting cone wall 8 to the intermediate casing 4; on the other hand, the No. 1 bearing seat and the supporting conical wall 8 are allowed to perform relative displacement along the radial direction within a limited range, so that the connection between the No. 1 bearing and the intermediate casing 4 is partially decoupled, and the supporting rigidity of the fan rotor 3 at the No. 1 bearing is reduced. The reduction of the bearing stiffness at bearing number 1 enables to reduce the critical speed of the fan rotor 3, making it lower than the working speed and higher than the fan speed, exerting the rotor self-centering effect and reducing the FBO load transmitted to the intermediary casing 4.

Although the present invention has been described with reference to the preferred embodiments, it is not intended to limit the present invention, and those skilled in the art can make modifications and variations without departing from the spirit and scope of the present invention.

Claims

1. An aircraft engine bearing support assembly for supporting a bearing (1) comprising a mounting block (20), a rigid support structure (21) and a shock absorbing assembly (22), the rigid support structure (21) being fixedly attached to the mounting block (20) and adapted to be connected to the bearing (1) to support the bearing (1);

characterized in that the shock absorbing component (22) comprises a first connecting ring (221), a second connecting ring (222) and a shock absorbing body (220) made of damping energy-absorbing material;

the first connecting ring (221) and the second connecting ring (222) are respectively and fixedly connected to the mounting seat (20) in a sealing manner, and are used for being respectively and fixedly connected with the bearing (1) in a sealing manner so as to support the bearing (1), and jointly define a closed annular cavity (2a) with the mounting seat (20) and an outer ring (12) of the bearing (1); the damping energy-absorbing material is used for filling the annular cavity (2a) to form the shock-absorbing body (220);

the first connection ring (221), the second connection ring (222) and the shock-absorbing body (220) are used for bearing the load between the mounting seat (20) and the bearing (1) to generate deformation, so that impact energy is absorbed.

2. The aircraft engine bearing assembly according to claim 1, wherein the first and second connecting rings (221, 222) are thin-walled metal plates.

3. The aero engine bearing assembly according to claim 1, wherein the first connection ring (221) has a first curved structure (221a) having a wave shape, the second connection ring (222) has a second curved structure (222a) having a wave shape, and the shock absorbing body (220) is snap fitted with the first curved structure (221a) and the second curved structure (222a), respectively.

4. The aircraft engine bearing assembly according to claim 1, wherein said rigid support structure (21) has a weakened portion (21a), said weakened portion (21a) being adapted to break upon receiving an overload force, thereby allowing said shock absorbing assembly (22) to solely support said bearing (1).

5. The aero engine bearing assembly according to claim 1 wherein said damping energy absorbing material is a metal rubber or foamed aluminum.

6. An aircraft engine comprising a bearing (1), characterized in that the aircraft engine (100) further comprises an aircraft engine bearing assembly (2) according to any one of claims 1 to 5,

the rigid support structure (21) and the shock-absorbing member (22) together support the outer race (12) of the bearing (1) before failure of the rigid support structure (21), and the shock-absorbing member (22) alone supports the outer race (12) after failure of the rigid support structure (21).

7. An aircraft engine according to claim 6, characterised in that the aircraft engine (100) further comprises an intermediate casing (4) and a fan shaft (5), the bearing (1) comprising an inner ring (10), rollers (11) and an outer ring (12);

the inner ring (10) is fixedly connected with the fan shaft (5), and the mounting seat (20) is fixedly connected to the intermediate casing (4); the mounting seat (20) surrounds the outer ring (12) and is coaxial (A-A) with the bearing (1), the rigid supporting structure (21) is connected with the outer ring (12), and the first connecting ring (221) and the second connecting ring (222) are respectively connected with the outer ring (12) in a sealing mode.

8. An aircraft engine according to claim 7, characterized in that said first connecting ring (221) and said second connecting ring (222) are each annular plates centred on said central axis (A-A) and axially spaced apart so as to form said annular cavity (2a) in the shape of a ring; the annular cavity (2a) is centered on the central axis (A-A).

9. An aircraft engine according to claim 6, characterised in that the bearing (1) is a front bearing of a fan shaft (5) of the aircraft engine.

10. An aircraft engine according to claim 8, characterised in that, with radial reference, the inner edge of the first connecting ring (221) is in sealing connection with the outer ring (12), the outer edge of the first connecting ring (221) being in sealing connection with the mounting seat (20); and the inner edge of the second connecting ring (222) is in sealing connection with the outer ring (12) by taking the radial direction as a reference, and the outer edge of the second connecting ring (222) is in sealing connection with the mounting seat (20).